Turbine coolant flow system

ABSTRACT

In a turbine engine compressed air is delivered through a diffuser to a burner section and then through a turbine section. A portion of the compressed air from said diffuser is removed by an annular center body manifold and directed by hollow struts inwardly to an annular passageway where the compressed air flow is taken to the forward part of the turbine section. The flow from the annular passageway is then directed into a passageway having directing vanes for imparting a desired velocity and direction to the existing flow so as to be compatible with the rotating turbine disk onto which it flows. This air is then directed to cool blades on that disk and also passed through that disk to be directed to turbine blades on the next disk. The invention herein described was made in the course of or under a contract with the Department of the Air Force.

[ July 311, 1974 TURBINE COOLANT FLOW SYSTEM [75] Inventors: Bruce R. Branstrom, Riviera Beach;

Frank W. Huber, Palm Beach Gardens, both of Fla.

[73] Assignee: United Aircraft Corporation, East Hartford, Conn.

[22] Filed: Oct. 30, 1973 [21] 'Appl. No.: 411,124

Related US. Application Data [63] Continuation of Ser. No. 32,687, April 28, 1970,

abandoned.

[52] 11.8. C1 60/39.66, 60/3907, 60/3931, 415/115, 415/144, 415/175 [51] Int. Cl. F02c 7/18 [58] Field of Search 60/3966, 39.07, 39.31, 60/3932; 415/165, 175, 176, 144, 145, 115

[56] References Cited UNITED STATES PATENTS 2,682,363 6/1954 Lombard 60/3907 2,812,898 ll/1957 Bucll 60/3966 2,986,231 5/1961 Hellstrom 60/3907 2,988,325 6/1961 Dawson 415/115 3,394,543 6/1968 Slattery 60/3907 3,565,545 2/1971 Bobo 415/175 3,632,223 1/1972 Hampton 415/144 3,635,586 l/l972 Kent 60/3966 FOREIGN PATENTS OR APPLICATIONS 1,231,999 4/1960 France 60/3966 Primary Examiner-Douglas Hart Attorney, Agent, or FirmJack N. McCarthy [5 7] ABSTRACT In a turbine engine compressed air is delivered through a diffuser to a burner section and then through a turbine section. A portion of the compressed air from said diffuser is removed by an annular center body manifold and directed by hollow struts inwardly to an annular passageway where the compressed air flow is taken to the forward part of the turbine section. The flow from the annular passageway is then directed into a passageway having directing vanes for imparting a desired velocity and direction to the existing flow so as to be compatible with the rotating turbine disk onto which it flows. This air is then directed to cool blades on that disk and also passed through that disk to be directed to turbine blades on the next disk. The invention herein described was made in the course of or under a contract with the Department of the Air Force.

4 Claims, 3 Drawing Figures PREMIER- 3.826.084

SHEET 2 0f 2 COOLING AIR VELOCITY RELATIVE TO DISK g gg mgmg ASK mm TURBINE COOLANT FLOW SYSTEM This is a continuation, of application Ser. No. 32,687, filed Apr. 28, 1970 now abandoned.

BACKGROUND OF THE INVENTION While cooling air has been directed to turbine blades for cooling, and compressor air has been used for this purpose, the arrangement set forth herein is a different device for doing so as will be later described.

SUMMARY OF THE INVENTION This turbine coolant flow system is arranged to provide low coolant temperatures with a minimum of expended horsepower to provide maximum engine efficiency. The coolant is expanded through nozzle means in a tangential plane paralleling the direction of turbine disk rotation.

This system provides reduction in the amount of coolant required, an increase in turbine horsepower, and a reduction in the first disk seal leakage.

The requirement for coolant is reduced due to reduction in cooling air temperature. The turbine horsepower is increased since the disk does not have to pump the coolant up to wheel speed to have it flow through the holes in the disk. First disk seal leakage is reduced because the pressure ratio across the seal is reduced.

The preswirl nozzle means accelerates the air flow and directs it in a tangential direction approximately paralleling the rotation of the first disk permitting the lower temperature and pressure. Minimizing cooling and leakage air quantities by providing cooler air at lower pressures allows more air to remain in the engine main stream.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a sectional view taken through an engine casing showing the diffuser section, burner section, and turbine section incorporating the turbine coolant flow system.

FIG. 2 is an enlarged view taken along the line 2-2 of FIG. 1 showing the directional vanes which direct coolant flow on to the first turbine disk.

FIG. 3 is a velocity diagram showing the relative velocity of the air leaving the vanes and of the disk.

DESCRIPTION OF THE PREFERRED EMBODIMENT As shown in FIG. 1 an annular diffuser 2 is shown discharging into an annular burner section 4 which contains a plurality of burners 6 therearound. Compressor means (not shown) discharges into the annular inlet of the diffuser 2, this can be from a compressor section as shown in US. Pat. No. 3,301,526. The outer wall 8 of the annular burner section is a continuation of the outer wall of the diffuser 2. Pin means 10 are provided at a plurality of locations around the engine where the outer diffuser wall and burner wall 8 meet to fix the forward end of each burner 6 in position. An arm 12 extends forwardly from each burner having opening means through which a pin is positioned which is fixed to the outer burner wall. The inner wall 14 of the burner section is a continuation of the inner wall of the diffuser 2. This inner wall 14 has a cooperating wall 15 which forms an annular passageway 17 therewith extending from a point adjacent the outlet of the diffuser to the forward part of the turbine section 11. Spacer members 13 are used to properly space the walls 14 and 15.

The rearward end of the outer burner wall 8 is fixed to the outer wall 16 of the turbine section 11. A plurality of vanes 18 extend inwardly from the outer wall 16 and are connected at their inner ends by a composite annular member 20. The composite annular member 20 is connected to the rear end of the inner wall 14 by a solid wall member 28. The annular member 20 includes one labyrinth member 22 extending rearwardly as a cylindrical member and having inwardly extending lands for a purpose to be here-in-after described. An annular duct member 24 is positioned radially inwardly from cylindrical member 22 and extends axially having a passageway 26 therethrough. Passageway 26 has a plurality of vanes 27 therein which takes the air which is directed substantially axially thereto and turns it a predetermined amount for a purpose to be here-inafter described.

The forward part of the duct 24 is connected at its outer edge to member 20 by an outwardly extending flange and is connected at its inner edge to wall 15 by a solid wall member 30. The solid member 30 also has a second labyrinth member 32 extending rearwarclly at a position radially inwardly from the duct 24 as a cylindrical member and having inwardly extending lands for a purpose to be here-in-after described. Solid wall member 28 and solid wall member 30 form an annular connecting means between the rear end of passageway 17 and the inlet to passageway 26.

The rear end of each burner is supported by flange members 3 and 5 which extend forwardly from the forward part of the turbine section 11. These flange members are adapted to engage cooperating members fixed to the rearward part of each of the burner cans. While one means has been shown, the burner can may be fixed in position by any means desired.

Immediately downstream of vanes 18 is the first stage of a turbine rotor comprising a rotor disk 40 with blades 42 mounted thereon. The rotor disk 40 is mounted on a flange 41 which extends from a shaft 43 which is mounted for rotation within said engine. Said shaft being mounted for rotation and sealed where necessary by conventional means. The blades may be attached to the rotor by any means desired which permits a flow of fluid through the blade (as shown by the arrows). The rotor disk 40 has a forward side plate 44 fixed thereto and to the blades, and a rearward side plate 46 fixed thereto and to the blades. The forward side plate 44 has a circular flange means 48 extending forwardly therefrom with its free end being positioned between labyrinth member 22 and the outer wall of the annular duct member 24. This free end has a sealing surface on its outer side which mates with the lands on the member 22. A similar flange 49 extends forwardly from the rotor disk 40 with one side of its free end being positioned adjacent the lands of the labyrinth member 32 so as to provide a sealing action at that point. These flange members 48 and 49 are spaced from the annular member 24 and form a rotating annular passageway 50 therebetween.

A plurality of holes 52 extend from the passageway 50 between the connection of flange means 48 and 49 on the rotor disk 40 to the space formed between the side plates 44 and 46. Holes 54 extend from the passageway 50 between flanges 48 and 49 through the disk to a point on the rearward side of the disk. Openings (not shown) extend through the platform in each blade 42 to permit fluid flow from the space between the side plates 44 and 46 to the interior of the blades 42. Fluid in blades 42 is permitted to flow therefrom through openings 56 located in the trailing edge of the blades.

Downstream of the first stage of the turbine rotor is the second stage comprising a rotor disk 66 and blades 62. Disk 60 is also fixed to the flange 41. A web member 64 interconnects the outer edge of each of the rotor disks and has a plurality of lands extending outwardly therefrom to provide a labyrinth seal for sealing in a manner to be here-in-after described. An annular chamber 70 is formed between the disks 40 and 60 and web member 64. The rotor disk 60 has a forward side plate 66 fixed thereto and to the blades, and a rearward side plate 68 is fixed thereto and to the blades. A plurality of holes 72 extend between the chamber 70 and the space formed between the side plates 66 and 68. Openings (not shown) extend through the platform in each blade 62 to permit fluid flow from the space between the side plates 66 and 68 to the interior of the blades 62. Fluid in blades 62 is permitted to flow therefrom through openings located in the outer end of the blades.

Vanes 80 extend inwardly from the outer wall 16 of turbine section 11 between blades 42 and 62. Said vanes are fixed at their outer end and have a sealing member at their inner ends which cooperates with the lands on the web member 64. Vanes 90 are positioned downstream of the blades 62 and seal means 92 are provided between the side plates 68 and inner structure of the vanes 90.

Openings 94 are located in flange 41 so that fluid bleeding by labyrinth lands on member 32 can pass to the rear of turbine disk 60 and back into the flow path.

An annular manifold 100 is positioned around the forward part of the burner section 4 with an annular opening 101 directed forwardly so as to pick up air from the compressor discharge diffuser. This location provides the coolest and cleanest mid-span air for cooling. The annular manifold 100 has projections 102 extending from the top thereof, each with an opening, to receive the pin from pin means 10. The annular manifold has a plurality of thin hollow struts 104 to connect the annular manifold to the forward end of the annular passageway 17. An opening 105 in wall 14 opens the interior of each hollow strut into passageway 17.

It can now be seen that air leaving a compressor located forwardly of the diffuser 2 will flow through diffuser 2 into the burner section 4. A predetermined amount of compressor air is picked up by the annular manifold 100 and directed into annular passageway 17 by the plurality of hollow struts 104 through holes 105. The flow then passes through passageway 17 into the annular connecting means formed by solid wall member 28 and solid wall member 30 and directed to the inlet of passageway 26. The flow is then expanded through the vanes 27 and exits from the passageway 26 in a predetermined direction and at a desired velocity to be compatible with the rotation of the turbine rotor disk 40. A portion of flow then passes through holes 54 into annular chamber 70 and a portion flows through openings 52 into the area between side plates 44 and 46 where it then flows through the blades 42. The air in chamber 70 then flows through openings 72 to the area between side plates 66 and 68 and then passes through blades 62.

Calculations made for one construction showed that flow losses were low in the system and approximately 96 percent of compressor discharged total pressure was supplied to the inlet of annular duct member 24. The cooling air was then expanded through the full annulus of vanes 27 to a pressure equal to approximately 65 percent of compressor discharged total pressure. Actually the swirl velocity imparted to the cooling air is greater than the speed of the rotating passageway and work is then done on the rotating disk. In this expansion process, the gas temperature inside the chamber 50 dropped approximately 120F. This cooling air then progressed through the flow system. As stated here-in-before seal leakage is minimized because the disk pressure level has been dropped to approximately 65 percent of the compressor discharged total pressure.

We claim:

1. In combination, a turbine engine including an annular inlet section having a first outer wall and a first inner wall, an annular combustion section having a second outer wall connected to said first outer wall and a second inner wall connected to said first inner wall, a turbine section, said turbine section including a turbine rotor, said turbine rotor including a turbine disk mounted for rotation along with turbine blades mounted therearound, an annular manifold located immediately downstream of said annular inlet section, said annular manifold having an outer surface and an inner surface, said annular manifold having an annular inlet for receiving flow from said inlet section, said annular manifold being axially in line with the center of said annular inlet section with said annular inlet of said annular manifold being forwardly into the center of the annular inlet section for receiving flow passing therethrough, said annular combustion section having a third inner wall spaced from said second inner wall providing a first annular passageway, hollow struts extending between the inner surface of said annular manifold and the adjacent end of said first annular passageway,

a rotarting annular passageway extending forwardly from said disk, said first annular passageway having its downstream end located adjacent the rotating annular passageway so as to direct cooling air therefrom into the rotating annular passageway, passage means in said disk being in communication with said rotating passageway for directing flow to said blades for cooling.

2. A combination as set forth in claim 1 wherein projections extend from the outer surface of said annular manifold outwardly towards the second outer wall of the annular burner section, each projection having an opening extending into the end thereof, means being provided at a plurality of locations for extending through said second outer wall into the openings in said projections for supporting said annular manifold.

3. A combination as set forth in claim 2 wherein said means for extending into the openings are individually removable pins.

4. A combination as set forth in claim 1 wherein a burner is located in said annular combustion section, the rearward part of said burner being connected between said second inner wall and said second outer wall and positioned to exhaust into said turbine section, the forward end of said burner having an arm means projecting outwardly therefrom, said arm means having a nular manifold, said pin means extending through said second opening therethrough, said second opening second opening into its cooperating opening in a probeing aligned radially outwardly from an opening in a jection.

projection extending from the outer surface of said an- 33mg? "UNITED STATES PATENT OFFICE u CERTIFICATE OF CORRECTION Patent No. 3,326,0 Dated July 30,1974

Inventor(s) Bruce R. Branstrom and Frank W. H er It is certified that error appears in the above-identified patent and that said Letters Patent are hereby corrected as shown below:

Column 4, 1ine 37, "being" should be --facing-- Column 4, line 44, "rotarting" should be "rotating-- Signed and sealed this 26th day of November 1974.

(SEAL) Attest:

McCOY M. GIBSON JR. c. MARSHALL 1mm Attesting Officer Coumisaioner of Patentsv 

1. In combination, a turbine engine including an annular inlet section having a first outer wall and a first inner wall, an annular combustion section having a second outer wall connected to said first outer wall and a second inner wall connected to said first inner wall, a turbine section, said turbine section including a turbine rotor, said turbine rotor including a turbine disk mounted for rotation along with turbine blades mounted therearound, an annular manifold located immediately downstream of said annular inlet section, said annular manifold having an outer surface and an inner surface, said annular manifold having an annular inlet for receiving flow from said inlet section, said annular manifold being axially in line with the center of said annular inlet section with said annular inlet of said annular manifold being forwardly into the center of the annular inlet section for receiving flow passing therethrough, said annular combustion section having a third inner wall spaced from said second inner wall providing a first annular passageway, hollow struts extending between the inner surface of said annular manifold and the adjacent end of said first annular passageway, a rotarting annular passageway extending forwardly from said disk, said first annular passageway having its downstream end located adjacent the rotating annular passageway so as to direct cooling air therefrom into the rotating annular passageway, passage means in said disk being in communication with said rotating passageway for directing flow to said blades for cooling.
 2. A combination as set forth in claim 1 wherein projections extend from the outer surface of said annular manifold outwardly towards the second outer wall of the annular burner section, each projection having an opening extending into the end thereof, means being provided at a plurality of locations for extending through said second outer wall into the openings in said projections for supporting said annular manifold.
 3. A combination as set forth in claim 2 wherein said means for extending into the openings are individually removable pins.
 4. A combination as set forth in claim 1 wherein a burner is located in said annular combustion section, the rearward part of said burner being connected between said second inner wall and said second outer wall and positioned to exhaust into said turbine section, the forward end of said burner having an arm means projecting outwardly therefrom, said arm means having a second opening therethrough, said second opening being aligned radially outwardly from an opening in a projection extending from the outer surface of said annular manifold, said pin means extending through said second opening into its cooperating opening in a projection. 